A Review of Some Current Problems in Gas Turbine Secondary Systems

1997 ◽  
Author(s):  
A. B. Turner ◽  
C. A. Long ◽  
P. R. N. Childs ◽  
N. J. Hills ◽  
J. A. Millward
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Gas Turbine ◽  
Flow Distribution ◽  
Rotating Shaft ◽  
High Pressure Compressor ◽  
Secondary Systems ◽  
Secondary Air ◽  
Cooling Air ◽  
Pressure Compressor

This paper reviews the current position of five major problem areas in gas turbine secondary air system design. Although the problems are of primary interest to the designer of the coolant flow paths, since they directly affect the temperature, the stresses and thus the life of the major rotating components, three of the problems interact with the main gas path and are thus also the concern of the mainstream aerodynamicist. The five problems reviewed are: prediction of the flow distribution and heat transfer in the high pressure compressor drive cone cavity from the turbine to the rim of the HP compressor running underneath the combustion chamber, the flow penetration and heat transfer in the multiple rotating cavities formed by the multiple discs of the high pressure compressor with a rotating shaft running through the bores; the control of ingestion of hot turbine mainstream gas into the rotor-stator wheelspaces through the rim-seals; the problem of compressor and turbine stator-well heating, particularly compressor stator-wells in which excessive temperatures have been occasionally measured and finally, the pre-swirl coolant system which has to take the blade cooling air across from the stationary casing to the rotating turbine disc in the most advantageous manner.

Part Load Conditions of Complex Cycle Power Plants With Intercooled Gas Turbine

1996 ◽  
Author(s):  
A. Peretto
Keyword(s):  
High Pressure ◽  
Gas Turbine ◽  
Working Conditions ◽  
Power Plants ◽  
Pressure Level ◽  
Combined Cycle ◽  
Brayton Cycle ◽  
Part Load ◽  
High Pressure Compressor ◽  
Pressure Compressor

The present paper evaluates the behavior, in design and part load working conditions, of a complex gas turbine cycle with multiple intercooled compression, and the optional preheating of the air at the high pressure compressor outlet by means of the gas turbine outlet hot gas. The results are then compared with those obtained by a Brayton cycle gas turbine, with or without preheating of the air at the high pressure compressor outlet. Subsequently, the performance of complex combined cycles, with intercooled gas turbine as topper and one, two or three pressure level steam cycle as bottomer, in design and part load working conditions is also evaluated. The performance of these complex combined plants is then compared with that obtained by a Brayton cycle gas turbine as topper and one, two or three pressure level steam cycle as bottomer. Part load working conditions are realized by varying either the inlet guide vane angle of the first compressor nozzles or the maximum temperature at the combustor outlet. The study shows that in part load working conditions obtained by varying IGV, the complex cycles, in the examined gas turbine or in the combined cycle power plants, give conversion efficiencies decidedly greater than those obtainable by varying combustor exit temperature. Furthermore it is found that these complex power plant efficiencies, in part load working conditions, are far greater than those obtained by the Brayton cycle gas turbine, or by combined cycle with Brayton cycle gas turbine as topper, if IGV adjustment is adopted. If power variation is obtained with combustor outlet temperature adjustment, the efficiencies of the combined power plants with complex or Brayton cycle gas turbines, are substantially the same, for the same relative power variation.

Theory and Application of Axisymmetric Endwall Contouring for Compressors

2011 ◽  
Author(s):  
Georg Kro¨ger ◽  
Christian Voß ◽  
Eberhard Nicke ◽  
Christian Cornelius
Keyword(s):  
High Pressure ◽  
Pressure Gradient ◽  
Gas Turbine ◽  
Static Pressure ◽  
Aerodynamic Performance ◽  
Leakage Flow ◽  
High Pressure Compressor ◽  
Blade Tip ◽  
Stage Efficiency ◽  
Pressure Compressor

Engine operating range and efficiency are of increasing importance in modern compressor design for heavy duty gas turbines and aircraft engines. These highly challenging objectives can only be met if all components provide high aerodynamic performance and stability. The aerodynamic losses of highly loaded axial compressors are mainly influenced by the leakage flow through clearance gaps. Especially the leakage flow due to the radial clearances of rotor blades affects negatively both, the efficiency and the operating range of the engine. Recent publications showed that the clearance flow and the clearance vortex can be influenced by an additional static pressure gradient at the outer casing, which is created by an axisymmetric wavy casing shape. A notable performance increase of up to 0.4% stage efficiency at design point conditions was reported for high pressure stages with large tip clearance heights [1] as well as for a transonic stage with a relatively small radial clearance gap [2]. An analytic approach to predict the effects of axisymmetric casing contouring has been developed at DLR, Institute of Propulsion Technology, and is outlined in the first part of this work. The characteristic behavior of the clearance vortex in an adverse pressure gradient is discussed by means of an inviscid vortex model [3]. The critical vortex parameters are isolated and related to the static pressure increase due to the casing contour. The second part illustrates the application of an axisymmetric endwall contour. A three dimensional optimization of the outer casing and the corresponding blade tip airfoil section of a typical gas turbine high pressure compressor stage with a high number of free variables is presented. The optimization led to a significant increase in aerodynamic performance of about 0.8% stage efficiency and to a notable reduction of the endwall blockage at ADP conditions. Furthermore, an improved off-design performance was found and a simple design rule is given to transfer both, the casing contour and the blade tip section modification on similar high pressure compressor blades. Based on these design rules the results of the optimized stages were applied to the rear stages of a Siemens gas turbine compressor CFD model. An increase of 0.3% full compressor performance was reached at design point conditions.

Measurement and Computation of Heat Transfer in High-Pressure Compressor Drum Geometries With Axial Throughflow

1997 ◽  
Vol 119 (1) ◽  
pp. 51-60 ◽  
Author(s):  
C. A. Long ◽  
A. P. Morse ◽  
P. G. Tucker
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Stokes Equations ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Multigrid Algorithm ◽  
Numerical Predictions ◽  
High Pressure Compressor ◽  
Dependent Form ◽  
Pressure Compressor

This paper makes comparisons between CFD computations and experimental measurements of heat transfer for the axial throughflow of cooling air in a high-pressure compressor spool rig and a plane cavity rig. The heat transfer measurements are produced using fluxmeters and by the conduction solution method from surface temperature measurements. Numerical predictions are made by solving the Navier–Stokes equations in a full three-dimensional, time-dependent form using the finite-volume method. Convergence is accelerated using a multigrid algorithm and turbulence modeled using a simple mixing length formulation. Notwithstanding systematic differences between the measurements and the computations, the level of agreement can be regarded as promising in view of the acknowledged uncertainties in the experimental data, the limitations of the turbulence model and, perhaps more importantly, the modest grid densities used for the computations.

Corrosion Studies of High and Low Pressure Compressor Blades of a Gas Turbine

2005 ◽  
Author(s):  
A. Boschetti ◽  
E. Y. Kawachi ◽  
M. A. S. Oliveira
Keyword(s):  
High Pressure ◽  
Gas Turbine ◽  
Low Pressure ◽  
Compressor Blade ◽  
Compressor Blades ◽  
X Ray ◽  
Corrosion Studies ◽  
High Pressure Compressor ◽  
Pressure Compressor ◽  
Base Superalloy

This work presents preliminary results of corrosion studies for three blades, one of the low pressure compressor and two of two different stages of the high pressure compressor of a gas turbine, which has been operating for 5,000 hours. Scanning Electron Microscopy (SEM), Energy Dispersive X-ray Spectroscopy (EDS), X-ray diffraction (XRD), Electrochemical Impedance Spectroscopy (EIS) in aqueous solution containing chloride, and Atomic Absorption Spectrometry (AAS) were used to characterize the blades surfaces. The SEM and EDS results showed that the homogeneity and amount of contaminants, such as sodium, potassium, calcium, magnesium, chloride and sulphur are bigger in the high pressure compressor blade surfaces than in the low pressure compressor blade surface. The EIS results showed that the degradation process in turbine compressor blades increases with the temperature and pressure increase inside the compressors and depends of the blade composition. The low pressure compressor blade, which was made of a Ti base superalloy exhibited smaller corrosion resistance (smallest charge transfer resistance value (Rct)) than the two high pressure compressor blades, which were made of a Fe base superalloy. However, despite of its lower resistance to corrosion, after 5,000 hours of service, the low pressure compressor blade did not present pitting corrosion while the high pressure compressor blades did.

Experimental Investigation of Mistuning Effects on High Pressure Compressor Stators

2009 ◽  
Author(s):  
Chao Zhang ◽  
Aldo Abate ◽  
John Crockett ◽  
Eric Ho
Keyword(s):  
High Pressure ◽  
Distribution Pattern ◽  
Gas Turbine ◽  
Strain Gauge ◽  
Experimental Results ◽  
Turbine Engines ◽  
Gas Turbine Engines ◽  
High Pressure Compressor ◽  
Circumferential Distribution ◽  
Pressure Compressor

High pressure compressor (HPC) stator vanes of small gas turbine engines frequently have high circumferential variation of vibratory stress. This is very important for vibratory stress measurement by strain gauge tests and structural high cyclic fatigue characterization. The current paper presents experimental results of studying the effects of HPC stator angular positions in gas turbine engines on the circumferential distribution pattern of vibratory stress. Strain gauge tests were done on a stator with cantilevered vanes. Each vane had a strain gauge deployed at the same location. The stator was installed in gas turbine engines at two different angular positions during strain gauge tests. The experimental results show that more than one resonant peak occurred for a given vibratory mode and engine order resonance. The frequencies of resonant peaks were close to one another. The circumferential distribution of maximum vibratory stress (i.e., the maximum magnitude of these resonant peaks) with respect to the stator itself has a similar pattern at the two different angular positions. This clearly indicates the distribution pattern does not follow the gas-path aerodynamic pressure, but follows the stator angular positions. The frequency of the maximum vibratory stress was found to vary from sector to sector instead of from vane to vane; the vanes in each sector have a same frequency. Mistuning analysis was performed on the HPC stator to illustrate a number of resonant peaks and the sector-to-sector frequency variation of the maximum vibratory stress. The approach of “subset of nominal system modes” (SNM) [1, 2] was employed for mistuning analysis and the frequency distributions of stator vanes obtained by bench frequency response tests were used as input data. At the end, one might conclude that the high circumferential variation of vibratory stress be related to mistuning effects due to small variations in vane properties.

Analysis on Turbine Cooling Air Bleeding of Intercooled Recuperated Turbofan Engine

2015 ◽  
Author(s):  
Hao Gong ◽  
Zhanxue Wang ◽  
Li Zhou ◽  
Xiaobo Zhang ◽  
Jingkai Wang
Keyword(s):  
High Pressure ◽  
Combustion Chamber ◽  
High Pressure Turbine ◽  
Turbofan Engine ◽  
Turbine Cooling ◽  
High Pressure Compressor ◽  
Cooling Technology ◽  
Technology Level ◽  
Cooling Air ◽  
Pressure Compressor

In order to further improve the intercooled recuperated turbofan engine (IRT) performance, the possible high pressure turbine (HPT) cooling air bleeding schemes were analyzed. There are two HPT cooling air extraction sections, i.e. the high pressure compressor exit (forward to the recuperator cold section inlet) and the combustion chamber inlet (back from the recuperator cold section outlet). The analysis results indicate that, bleeding the HPT cooling air from the combustion chamber inlet has the potential to reduce the engine specific fuel consumption. And to determine the most suitable HPT cooling air bleeding scheme, effects of allowable turbine blade metal temperature, turbine cooling technology level, engine weight addition, different intercooler and recuperator effectiveness should be taken into account.

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